1. Field of the Invention
The present invention relates generally to the field of fabricating advanced composite aerostructure articles and more particularly, but not by way of limitation, to a method for fabricating an advanced composite aerostructure article having an integral co-cured fly away hollow mandrel.
2. Prior Art
There is a growing trend in the aerospace industry to expand the use of advanced composite materials for a diverse array of structural and dynamic aerostructural applications because of the strength-to-weight advantage provided by composite materials. One particular application for the use of such advanced composite materials lies in the fabrication of advanced composite articles such as panels for nacelles for aircraft jet engine propulsion systems. Such structural articles as fan cowls generally comprises inner and outer composite skins, which are formed from composite materials such as graphite or an aromatic polyamide fiber of high tensile strength that are embedded in a resinous matrix, e.g., epoxy, having a honeycomb core material interposed therebetween in the instance of a fuselage panel.
In the instance of an aerospace article such as a fan cowl, one or more stiffening members are affixed to the outer skin and covered with an inner skin for efficiently transmitting and/or reacting axial and/or bending loads to which the fan cowl is subjected.
There are two techniques currently employed for bonding through autoclave processing a composite stiffening member in combination with a composite structural panel: (1) the co-cured bonding method and (2) the secondary bonding method. Both methods are disadvantageous in requiring costly non-recurring tooling and/or costly recurring manufacturing steps.
A typical composite sandwich panel intended for use as an aerostructure article is normally fabricated using two autoclave cured inner and outer composite skins that are formed by using a curing cycle with heat, pressure, and a unique tool for each skin. A sandwich panel is then made up using a composite bond jig, tool or fixture with the pre-cured face skin laid-up on the bond jig tool followed by a ply of film adhesive, a honeycomb aluminum or non-metallic core of a given thickness, another ply of film adhesive and finally the previously pre-cured inner skin. The bond jig that is used to fabricate the sandwich panel is generally the same tool that was used to create the outer composite skin. A plurality of closure plies of uncured composite material are layed up and the assembled sandwich panel are cured during their final assembly stage. This sandwich panel is then vacuum bagged to the composite bond jig and again cured in an autoclave under high pressure and heat.
Thus, at least three very expensive and man-hour consuming cure cycles have gone into the fabrication of this exceptionally strong but lightweight composite/honeycomb core sandwich panel. At least two different and expensive tools are needed in this process. Manufacturing flow time is very long, energy use is high and the manufacturing floor space required is excessive.
The co-curing method envisions curing the composite inner and outer skins that are laid-up with a layer of adhesive film and honeycomb core in one cure cycle in the autoclave. A co-cured panel is desirable in that it is less expensive to fabricate since only one bond jig tool is required, only one cure cycle is needed, it is less labor intensive, it requires less floor space to accomplish, and a much shorter manufacturing flow time is achieved. However, co-curing an aerostructure panel has never achieved wide spread acceptance because of the large loss of panel strength and integrity that is lost due to the lack of compaction of the composite plies placed over and under the honeycomb core details. The composite plies dimple into the center of each core cell with nothing but the cell walls to compact the composite skins. The only way to overcome this knockdown factor is to add extra plies which creates both unwanted weight and excess cost. Thus, because of these constraints co-cured aerostructure panels are not widely manufactured in the aerospace industry.
There are other particular problems when a honeycomb core element is used to provide a stiffening element for an aerospace article such as a fan cowl. As Hartz et al described in U.S. Pat. No. 5,604,010 concerning a xe2x80x9cComposite Honeycomb sandwich Structurexe2x80x9d, with a high flow resin system, large amounts of resin can flow into the core during the autoclave processing cycle. Such flow robs resin from the laminate, introduces a weight penalty in the panel to achieve the desired performance, and forces over-design of the laminate plies to account for the flow losses. To achieve the designed performance and the corresponding laminate thickness, additional plies are necessary with resulting cost and weight penalties. Because the weight penalty is severe in terms of the impact on vehicle performance and costly in modern aircraft and because the flow is a relatively unpredictable and uncontrolled process aerospace design and manufacture dictates that flow into the core be eliminated or significantly reduced. In addition to the weight penalty from resin flow to the core, it was discovered that microcracking that originated in the migrated resin could propagate to the bond line and degrade mechanical performance. Such microcracking potential has a catastrophic threat to the integrity of the panel and dictates that flow be eliminated or, at least, controlled.
Unfortunately, the use of honeycomb core as a stiffener for elements in a aerostructure article such as a fan cowl, or in a structural panel has other deleterious effects, two of the greatest drawbacks to aluminum core being its inherent significant cost and corrosion. To minimize galvanic corrosion of the core caused by contact with the face skins, isolating sheets are interposed between such aluminum core and the face skins. Also, the aluminum core has an inherent cost and also must be machined to a desired shape in an expensive process. The honeycomb core may also be subject to crush during manufacture and thereby limits the pressures that may be used in autoclave processing. Also, the honeycomb core if damaged in use has a spring back effect which makes the detection of such damage more difficult. Thus, the processing of an aerospace advanced composite article is limited to an autoclave pressure of not greater than 35 psi rather than an advanced pressure that would increase the strength of the resultant advanced composite article.
In providing reinforcing mandrels for stiffener elements, such as hat sections, for aerospace advanced composite structural panels it is also known to provide a composite stiffening member in the form of a polyimidide foam mandrel fabricated by machining a core mandrel to a desired shape. Obviously, the machining of the core mandrel is expensive and time consuming and further introduces the problem of properly bonding the core mandrel to inner and outer skins.
Therefore, a great need has arisen for a practical method of readily producing stiffened, fiber-reinforced composite structures useful in the construction of integrally stiffened components for aerospace applications which are cost and labor efficient and which save time in the fabrication process.
Accordingly, it is an object of the present invention to provide a method for fabricating aerostructure advanced composite articles that eliminates honeycomb core in stiffening elements, provides a lighter weight assembly and is easier to repair.
Another object of the present invention is to reduce the layup cost of known advanced composite co-cure assemblies by at least 15% and to increase assembly strength over such existing co-cure assembly by being able to utilize advanced pressures in autoclave processing.
Yet another object of the present invention is to improve the quality of assembly of such co-cured advance composite assemblies and thereby increase customer satisfaction.
A further object of the present invention is to provide a process that provides an assembly that can be manufactured in one manufacturing cell from raw material to final product.
Yet further objects of the present invention are to reduce the cost of post bond and final assembly work for the final co-cured assembly, which assembly will readily indicate damage to an improved stiffening element.
The foregoing has outlined some of the more pertinent objects of the invention. These objects should be construed to be merely illustrative of some of the more prominent features and applications of the intended invention. Many other beneficial results can be attained by applying the disclosed invention in a different manner or by modifying the invention within the scope of the disclosure. Accordingly, other objects and a fuller understanding of the invention may be had by referring to the summary of the invention and the detailed description of the preferred embodiments in addition to the scope of the invention defined by the claims taken in conjunction with the accompanying drawings.
The foregoing problems are overcome and other advantages are provided by a new and improved method for fabricating an advanced composite aerostructure article from fiber reinforced composite material and incorporating a hollow stiffened graphite fabric mandrel that becomes an integral part of such article.
In accordance with a preferred embodiment of the present invention, a suitable lay-up mandrel or composite bond jig (COBJ) having a predetermined upper lay-up surface receives a first composite uncured layer to provide one surface of a composite aerostructure article, the first uncured layer having at least one uncured resin-impregnated laminate layer. A bond jig may be referred herein as a tool or fixture or lay up mandrel. A hollow mandrel having a predetermined cross sectional shape is layed up on and adhesively tacked to the upper surface of such first uncured composite layer which preferably forms the outer skin of an aerospace article. The hollow mandrel is preferably a spirally wound longitudinally extending shaped formed of a stiffened graphite fabric layer. A second uncured composite layer additional plies of uncured campsite layers is assembled by lay up over the upper outer surface of the bond jig and at least a portion of the first uncured composite layer, the second uncured layer having at least one uncured resin-impregnated layer and, preferably, a plurality of plies that does not cover the open ends of the hollow mandrel.
A suitable flexible tube bag is then positioned within the hollow mandrel and extends beyond its open ends. A suitable vacuum bag is then placed over the lay-up assembled on the layup mandrel or bond jig and removably secured to such lay-up assembly and to the ends of the tube bag as by a suitable sealant tape which permits the ends of the hollow mandrel to remain open. The vacuum bag is exhausted to secure the lay-up assembly upon the lay up mandrel. The lay up assembled on the lay-up mandrel is then subjected in a suitable autoclave to a cure cycle of predetermined pressure and temperature to cure such resin impregnated layers to form a unitary co-cured one piece aerostructure article having a reinforcing hollow hat section provided therein. Since the ends of the hollow mandrel remain open during the cure cycle the lay-up assembly may be subjected to an increased pressure to strengthen the resultant article. Following the cure cycle the bond jig is removed from the autoclave and the co-cured article is removed from the bag and trimmed as desired.
The cross sectional area of the hollow mandrel may be varied as desired to provide a preferred shape that may provide a cross sectional area that may be changed to accommodate other structures that may be in close vicinity to such article in its ultimate use.
To accommodate the need to further strengthen the aerostructure article in certain areas where the configuration of aerospace article or part does not allow access for a tube bag to reach the outer periphery of a hollow mandrel, a hollow mandrel containing a suitable tube bag may be secured to the surface of the first uncured skin layer and in contact with spaced longitudinally extending reinforcing hat sections prior to application of the subsequent composite layers. In this instance, the internal air bag is provided with a suitable vent means extending through the hollow mandrel, the uncured xe2x80x9chatxe2x80x9d plies of the second uncured layer and the primary vacuum bag to permit the interior of such bag to freely communicate with the atmosphere in the interior of the autoclave so that the pressure provided within such transversely provided hollow mandrel portion is the same as that pressure experienced on the exterior surface of the lay-up assembly.
A suitable tooling means such as a xe2x80x9cthrow in blockxe2x80x9d may be inserted in an end of a hollow mandrel during lay up, the tooling block being configured to allow installation of a tube bag through such block and the hollow mandrel and to provide a particular configuration for such end during the curing process. Thus, a suitable mechanical attachment such as a hinge or latch member may be easily secured to the end of the co-cured hat section without the necessity of providing shimming members as is now required for honeycomb core hat sections for advanced composite aerostructure articles.
The foregoing has outlined rather broadly the more pertinent and important features of the present invention in order the detailed description of the invention that follows may be better understood so that the present contribution to the art may be more fully appreciated. Additional features of the invention will be described hereinafter which form the subject of the claims of the invention. It should be appreciated by those skilled in the art that the conception and the disclosed specific embodiment may be readily utilized as a basis for modifying or designing other structures and methods for carrying out the same purposes of the present invention. It should also be realized by those skilled in the art that such equivalent constructions and methods do not depart from the spirit and scope of the invention as set forth in the happened claims.